This invention relates generally to the field of composite materials and more particularly to a method and apparatus for fabricating composite structures.
Many applications today call for the fabrication of components from xe2x80x9ccompositexe2x80x9d materials, also known as fiber-reinforced plastics. Fiber-reinforced plastics are comprised of reinforcing fibers that are positioned in a polymer matrix. Commonly, the reinforcing fibers are fiberglass, although high strength fibers such as aramid and carbon are used in advanced applications, such as aerospace applications. The polymer matrix is a thermoset resin, typically polyester, vinyl ester, or epoxy. Specialized resins, such as, phenolic, polyurethane and silicone are used for specific applications.
Composite materials may be formed using numerous fabrication process. One such process that is common in the aerospace industry is a lay up process. In a typically lay up process, layers of reinforcing fiber are laid in a mold by hand or by a placement machine. Liquid resin is then poured on the fiber materials such that the resin fills the spaces between the fibers. The materials may then be cured at room temperature or in an autoclave and the liquid resin turns into a solid thermoset. The fibers are thus impregnated in the solid thermoset resin and reinforce the resin. Alternatively, layers of fibers can be pre-impregnated with resin and then partially cured to form layers of xe2x80x9cprepregxe2x80x9d material. After this partial curing, the resin has not completely set, and the prepreg layers are flexible and can be shaped in or around a mold or forming tool. Once the prepreg layers are so shaped, the prepreg is then completely cured in an autoclave to form a fiber-reinforced laminate.
Composite structures often have several discrete composite components. For example, composite aircraft structure includes a composite skin, a number of stiffening members and other support structures. In metal aircraft wings, the various parts are formed separately and then fastened together using fastening methods such as welding and riveting. However, due to the nature of composite components, such fastening methods are not useful. Instead, the composite components are typically formed together using methods such as lay ups and autoclave curing, described above.
Due to the nature of forming tools that are traditionally used, these components are typically fabricated in a sequential fashion. For example, in the fabrication of a composite wing, the skin is first molded and cured. One set of stiffening members is then formed on the wing skin, and the entire structure is cured again. Further stiffening members and other structural supports are then sequentially added to the structure and cured. This method of fabrication is extremely expensive and time-consuming, and greatly increases the costs of advanced composite structures such as aircraft components.
Therefore, a need has arisen for a new method and apparatus for fabricating composite structures that overcomes the disadvantages and deficiencies of the prior art.
A method for fabricating composite structures is disclosed. The method includes providing a skin that includes one or more layers of uncured composite material, and providing a flexible hinge tool that has first and second tooling portions coupled with a flexible hinge. The first and second tooling portions of the hinge tool each have a molding surface, and the tooling portions are configured to form at least a portion of a stiffening member. The method further includes laying up at least one layer of uncured composite material on the molding surfaces to form the stiffening member.
Furthermore, the flexible hinge tool is positioned on the skin such that the uncured composite material laid up on the first tooling portion of the hinge tool contacts the skin, and such that the uncured composite material on the second tooling portion extends from the skin to form an upstanding segment of the stiffening member. The skin and the uncured composite material on the flexible hinge tool are simultaneously cured to form a stiffened composite structure. The method also includes the step of removing the flexible hinge tool by bending the tool at the flexible hinge such that the first portion of the tool peels away from the skin, and such that the second portion of the tool peels away from the upstanding segment of the stiffening member.
In another embodiment, an apparatus for fabricating composite structures is provided. The apparatus includes first and second tooling portions that are configured to support at least one layer of composite material during curing to form a composite structure. The apparatus further includes a flexible hinge that is disposed between and couples the first and second tooling portions. The flexible hinge couples the tooling portions such that they may be peeled away from the layers of composite material after curing by bending the apparatus at the flexible hinge.
A technical advantage of the present invention is that a method for fabricating composite parts is provided. Another technical advantage is that this method allows most, if not all, of the components of a stiffened composite skin to be fabricated using a single curing process, thus reducing time and expense. A further technical advantage is that the present invention allows stiffening members to be positioned closely together on a composite structure. The above advantages are due, in part, to the ability of flexible hinge tools incorporating teachings of the present invention to be easily removed from between closely spaced components. Another technical advantage is the ability of such flexible hinge tools to have numerous configurations for tooling various stiffening members and other composite parts.